This invention relates generally to supersonic aircraft engine air inlet and nozzle systems, and more particularly to enhancement of efficiency of such systems. It also relates to reducing or eliminating the requirement for stabilizing bleed air.
Supersonic aircraft engine air inlet systems are faced with a difficult challenge in maximizing performance of the aircraft. At supersonic speeds the engine inlet must slow the air velocity to less than the speed of sound, typically less than Mach 0.6 at the engine inlet face. To accomplish this, the inlet must subject the air to a shock system. In passing through the shock system losses in total pressure occur which reduce the net thrust and net thermal efficiency of the engine. These losses can be reduced to negligibly low levels by incorporating a suitably shaped isentropic compression surface, however as the flow is decelerated near Mach one, inlet stability problems occur for such high efficiency inlets as flow approaches two possible flow conditions. These are called subcritical where the flow is subsonic ahead of the inlet throat (the point of minimum cross-sectional normal to the local flow) or supercritical where the flow passes the throat supersonically with a series of oblique shock waves.
Inlets are typically designed to place a final terminal shock of a given strength near the throat where the flow will pass from supersonic to subsonic flow, the strength of which is a measure of relative flow stability. A very weak terminal shock, for example decelerating the flow from Mach 1.1 down to Mach 0.91, will exhibit very little total pressure loss, but would be prone to flow instabilities such as “buzz” where the inlet rapidly oscillates from subcritical to supercritical operation. Such instabilities could be triggered by changes in temperature, moisture, or flow angle such as from gusts. To prevent this, supersonic inlets typically are designed to operate with a terminal shock strength between 1.2 to 1.3, which results in a small but un-recoverable loss in total pressure of 0.8 to 2%.
In addition, supersonic inlets are typically fitted with bleed air systems to remove a small portion of the boundary layer on the compression surface at the terminal shock location. The boundary layer bleed is needed to hold the shock at the design location, prevent instability, and to prevent boundary layer separation. This can be explained as follows: A shock system represents a very strong adverse pressure gradient to a boundary layer which will cause the boundary layer to thicken or separate. A rule of thumb is that a Mach 1.3 normal shock strength will induce separation of even a very fresh boundary layer. Even if not separated, the boundary layer will thicken at the shock, reducing the effective throat area. Reducing the throat area in turn strengthens the shock, further increasing the adverse pressure gradient and reducing the effective throat area, and so forth. The result can either be “buzz” or the shock may move forward to a point of a stronger terminal shock well ahead of the intended location. This condition results in significantly higher overall pressure losses and variable pressures to the engine (distortion)
The stabilizing bleed system represents an additional loss in net thrust of the system, as it requires added pressure loss (or mechanical pumping) to induce the bleed flow.
A further consequence of low loss nearly isentropic compression for external compression inlets is cowl wave drag. In order to generate the shocks for low loss supersonic compression the flow must be turned from the free stream direction. The greater the required efficiency or design Mach number, the greater the flow turning angle. For a typical external compression inlet with some spillage around the inlet lip (local mach/mach=1 or M/M*<1) the flow spilling around the outside of the inlet lip incurs a drag penalty (additive drag). The additive drag is a function of the flow angle, and thus the total net thrust becomes a trade-off, between pressure recovery loss through the engine inlet compression system and inlet additive drag. The maximum thrust occurs with less than isentropic compression (see AIAA 2004-4492 “Multidisciplinary Optimization of a Supersonic Inlet Using a Cartesian CDF Method” paper by Rodriguez).
Present day commercial supersonic aircraft concepts anticipate the use of bypass fanjet engines rather than the traditional turbojets such as on Concorde. The bypass fanjet is distinct from the turbojet in bypassing additional air from the initial fan stages around the outside of the engine core, (compressor, combustor and turbine), providing improved propulsive efficiency and reduced noise. A characteristic of the fanjet engine is that reductions in net thrust from inlet pressure recovery losses are significantly lower for the outer fan air than for the inner core air destined to pass through the core of the engine.
The invention also relates generally to supersonic aircraft engine air inlet designs operating efficiently over a broad range of conditions from very low speeds for takeoff to very high speed cruise.
Jet powered aircraft derive thrust by means of turbojet or turbofan engines which induce flow through an air inlet, increase the pressure and temperature of the induced flow and exhaust it out an appropriate nozzle at higher velocity than it entered. A critical challenge for the successful design of supersonic aircraft is air inlet systems which can operate at low speed and high thrust conditions for takeoff and in flight conditions ranging from subsonic to transonic, and supersonic regimes. Typically an inlet designed for efficient low drag supersonic cruise features very thin sharp inlet lips. At the low speeds needed for takeoff and initial climb the engine requires a very high airflow and induces airflow velocities near the inlet lip much greater than the freestream velocity. This results in a “vena contracta” typical of flow through a sharp edged orifice which limits the flow volume and creates large flow separations, pressure losses and distortions which are unacceptable to the engine. An early solution to this dilemma was the “translating cowl” in which the inlet was made in two pieces such that the most forward portion incorporating the sharp supersonic lip moved forward away from rear portion of the inlet and exposed a second inlet suitable for ingesting additional air through the lateral opening created between the forward and aft inlet sections.
An additional challenge for supersonic inlets is accommodating the changing requirements with speed. Typically they incorporate a forward ramp or spike surface ahead of the enclosed portion of the inlet which presents an angle to the flow to generate a weak shock system to slow and compress the air before entering the enclosed portion of the inlet. The ideal ramp angle for such an inlet changes with Mach number.
A third difficulty is the changing characteristic of the airflow demands of the engine. Often as Mach number increases the engine will accept less air than provided by the inlet system, and the excess must be spilled around the inlet or bypassed through some auxiliary openings in the inlet internal and external surfaces. In supersonic flow it generally creates a smaller drag penalty on the aircraft to bypass air after it is taken into the inlet than to spill it ahead of the inlet. Many supersonic aircraft have incorporated complex and heavy variable ramp and bypass systems to accommodate these supersonic matching problems.
Improvements are needed to provide lighter, more efficient and less complex means for accommodating the diverse requirements of supersonic aircraft inlets.
The invention further relates generally to supersonic aircraft jet engine nozzle efficient integration with the aircraft fuselage, and also to engine nacelle efficient integration with the fuselage.
Jet powered aircraft derive the thrust required by means of engines which take in free-stream air, increase the pressure and temperature of the air, and reaccelerate that air to a higher velocity than when it entered. A critical part of the propulsion system is the nozzle, which takes the air which leaves the engine at high total pressure but reduced velocity and accelerates it to the higher exhaust velocity. For supersonic aircraft the pressure ratio (of engine exhaust total pressure divided by ambient pressure) exceeds the critical pressure ratio and requires an expansion of the exhaust from subsonic to supersonic velocity. The nozzle must provide a carefully designed flow path to allow this expansion with minimal loss in total pressure through shock waves. The flow path of a typical nozzle involves a decrease in area as flow is accelerated from subsonic velocity at the engine exhaust to a minimum throat area where the flow attains sonic velocity (Mach 1.0) and from there expands in area again to accelerate the flow to final supersonic velocity.
The most basic nozzle for such applications, is the convergent-divergent or C-D nozzle. The efficiency of the fixed C-D nozzle varies significantly with the different pressure ratios and operating conditions required of a supersonic aircraft, whereas it has been found that a “plug” nozzle provided comparable peak efficiency to a C-D nozzle with less efficiency loss away from the design operating condition. The plug nozzle consists of a circular outer cowl duct with an inner spike located in the center but projecting behind the exit plane of the outer duct. Most (but not necessarily all) of the supersonic expansion takes place on the externally exposed surface of the spike. Expanding a flow to supersonic speed with minimum pressure loss requires a nearly isentropic expansion and involves turning the flow through definite angle. Achieving maximum thrust from the nozzle requires that at its final accelerated velocity the flow must be approximately aligned with the flight direction. This in turn requires that prior to supersonic expansion the flow must be turned towards the spike, resulting in the external nacelle surface immediately ahead of the nozzle exit lip presenting a significant angle to the external flow. This angle forces the external flow to expand locally, creating a negative pressure zone and drag on the nacelle surface. This drag is termed “boat tail drag”.
There is need for improvements in jet engine nozzles that provide efficient thrust conversion over wide operating ranges. There is need for engine nacelle, fuselage and wing configurations in combinations that significantly reduce supersonic boat tail drag penalties.